A gas turbine engine includes a core engine having, in serial flow relationship, a high pressure compressor (also called a core compressor) to compress the airflow entering the core engine, a combustor in which a mixture of fuel and the compressed air is burned to generate a propulsive (combustion) gas flow, and a high pressure turbine which is rotated by the propulsive gas flow and which is connected by a larger diameter shaft to drive the high pressure compressor. A typical aircraft bypass turbofan gas turbine engine adds a low pressure turbine (located aft of the high pressure turbine) which is connected by a smaller diameter coaxial shaft to drive a front fan (located forward of the high pressure compressor) which may also drive a low pressure compressor (located between the front fan and the high pressure compressor). The low pressure compressor sometimes is called a booster compressor or simply a booster. It is understood that the term "compressor" includes, without limitation, high pressure compressors and low pressure compressors. A flow splitter, located between the fan and the first (usually the low pressure) compressor, separates the air which exits the fan into a core engine airflow and a surrounding bypass airflow. The bypass airflow from the fan exits the fan bypass duct to provide most of the engine thrust for the aircraft. Some of the engine thrust comes from the core engine airflow after it flows through the low and high pressure compressors to the combustor and is expanded through the high and low pressure turbines and accelerated out of the exhaust nozzle.
Cooling of engine hot section components, such as the combustor, is necessary because of the thermal "redline" limitations of the materials used in the construction of such components. Typically such cooling of the combustor is accomplished by using a significant amount of air which exits the compressor. This cooling air bypasses the combustion chamber and is used to cool the combustor (e.g., the combustor liners) as well as to cool, for example, turbine components. The cooling air, after cooling the combustor (and turbine components), re-enters the gas path downstream of the combustor. Because this cooling air is not available inside the combustion chamber, the combustor has to operate at a higher fuel to air ratio which results in a higher combustor temperature in order to provide a desired turbine inlet temperature which is required for engine power and efficiency. However, the higher combustion chamber temperature generates more undesirable NOx emissions.
Gas turbine engine NOx emissions from operation on liquid fuels is at least partially the result of stoichiometric fuel-air ratio in the vicinity of the liquid fuel droplets. Above the pseudo-critical pressure of jet fuel (approximately 350 psia), liquid droplets absorb heat by convection and radiation. Once the temperature of the droplet reaches the pseudocritical temperature (approximately 750.degree. F.), the droplet loses surface tension and disperses via air shear forces into a fuel-rich concentration of gaseous fuel and air molecules. With additional mixing with air, the gas mixture reaches ignition fuel-air ratio (still fuel rich) and bums. Since the engine overall fuel air ratio is fuel lean, there is a region where the gaseous fuel-air ratio is at or near stoichiometric concentration and the flame temperature is at a maximum (about 4000.degree. F.).
NOx is produced by the reactions N2+O-&gt;NO+N followed by O2+N-&gt;NO+0. The rate of the second reaction (O2+N) increases NOx by a factor of five over a peak combustor temperature of 3000.degree. F. when the fuel droplets burn at 4000.degree. F. The relative NOx increase is much greater for lower engine power levels.
It would be desirable to lower the NOx emissions of a gas turbine engine without adversely affecting the engine efficiency. It also would be desirable to achieve such reduced NOx emissions without significantly increasing the cost of the engine.